| Autor: |
Eric Besnard, Robert J. Shinavski, Andrea Wilson, Christopher Bostwick |
| Rok vydání: |
2009 |
| Předmět: |
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| Zdroj: |
45th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit. |
| DOI: |
10.2514/6.2009-5478 |
| Popis: |
Among the various approaches used for cooling the walls of a combustion chamber, ablative cooling is a simple and reliable solution. In that case, in order to reduce changes in throat geometry, an insert is usually used in that region. A proposed alternative is to use advanced ceramic matrix composite liners which are inserted throughout the thrust chamber in order to allow higher wall temperatures than ablatives with greatly reduced erosion rates. This paper details the design of a proof-of-concept 2200 N (500 lbf) thrust LOX/ethanol rocket engine. A Cf/SiCm ceramic matrix composite is used as liner in the chamber and throat for maintaining the combustion chamber internal dimensions. The remainder of the nozzle is made from a typical silica-phenolic ablative. Film cooling is added to maintain the C/SiC liner below 1800⁰C. The aluminum injector features a row of 24 split triplets and 24 film cooling orifices. A successful static fire test of the engine was conducted in November 2008, demonstrating a total burn duration of 33 seconds without significant erosion of the liner while significant ablation was noticed at the liner/ablative transition in the nozzle. The test demonstrated the viability of the concept for improving performance of ablative rocket engines. Future studies will concentrate on transitioning to LOX/methane propellants and reducing required ablative thickness and film cooling. This approach has the potential for improved performance of future ablative engines, such as the lunar ascent main engine now in development for NASA. |
| Databáze: |
OpenAIRE |
| Externí odkaz: |
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